The present disclosure relates to a gas turbine engine and, more particularly, to a cooling arrangement therefor.
Gas turbine engines, such as those that power modern commercial and military aircraft, generally include a compressor section to pressurize an airflow, a combustor section to burn a hydrocarbon fuel in the presence of the pressurized air, and a turbine section to extract energy from the resultant combustion gases.
The combustor section produces a circumferential temperature pattern referred to as a pattern factor that results in hot and cold streaks in the turbine section. Stationary components such as stationary vane arrays and Blade Outer Air Seals operate at the local pattern temperature and are thereby designed to withstand the local max temperature hot streaks which typically requires significantly higher dedicated cooling flow rates than rotational components such a blade arrays.